Rotorcraft control mode transition smoothing

ABSTRACT

In accordance with an embodiment, a method of operating a rotorcraft includes transitioning from a first mode to a second mode when a velocity of the rotorcraft exceeds a first velocity threshold. Transitioning between the first and second modes includes fading out a gain of a dynamic controller over a first period of time, and decreasing a value of an integrator of the dynamic controller over a second period of time. In the first mode, the translational speed of the rotorcraft is determined based on a pilot stick signal, and in the second mode, an output of an attitude rate controller is proportional to an amplitude of the pilot stick signal.

TECHNICAL FIELD

The present invention relates generally to a system and method for aflight control, and, in particular embodiments, to a system and methodfor control mode transition smoothing for a rotorcraft.

BACKGROUND

Fly-by-wire systems in aircraft, as opposed to mechanically controlledsystems, use electronic signals to control the flight surfaces andengines in the aircraft. For example, instead of having the pilotcontrols mechanically linked to the control surfaces via a hydraulicsystem, the pilot controls are electronically linked to a flightcomputer, which, in turn, controls flight surface actuators viaelectronic signals. By further interfacing the flight computer toaircraft sensors, sophisticated control algorithms may be used toprovide autopilot functionality, as well as to stabilize and control theaircraft.

While fly-by-wire systems have become commonplace in commercial andcivilian fixed wing aircraft, their adoption among rotorcraft, such ashelicopters, has been much slower due, in part, to the increasedcomplexity of controlling and stabilizing a rotorcraft. However, byadopting fly-by-wire systems in helicopters, safer operation may beachieved in difficult flight environments such as low speed, lowaltitude, degraded visual environments and inclement weather. Anotherarea in which fly-by-wire systems may benefit rotorcraft is in thereduction in pilot workload. By providing automated features such asstabilization in response to wind, control axis decoupling, positionhold and heading hold functionality, the pilot is freed up to focus onthe environment in which he flies.

One challenge in the design of fly-by-wire systems for rotorcraft istransitioning between different modes of operation that utilizedifferent control laws or algorithms. In some circumstances, the changein control algorithm may result in a physical transient during operationof the rotorcraft that might be physically discernable as a bump or joltby the pilot or passengers.

SUMMARY

In accordance with an embodiment, a method of operating a rotorcraftincludes transitioning from a first mode to a second mode when avelocity of the rotorcraft exceeds a first velocity threshold.Transitioning between the first and second modes includes fading out again of a dynamic controller over a first period of time, and decreasinga value of an integrator of the dynamic controller over a second periodof time. In the first mode, the translational speed of the rotorcraft isdetermined based on a pilot stick signal, and in the second mode, anoutput of an attitude rate controller is proportional to an amplitude ofthe pilot stick signal.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of the present invention, and theadvantages thereof, reference is now made to the following descriptionstaken in conjunction with the accompanying drawings, in which:

FIG. 1 illustrates an embodiment rotorcraft;

FIG. 2 illustrates a block diagram of an embodiment rotorcraft flightcontrol system;

FIG. 3 illustrates a block diagram of an embodiment flight controlsystem;

FIG. 4 illustrates a block diagram of a further embodiment flightcontrol system;

FIG. 5 illustrates a block diagram of an embodiment TRC mode attitudecontroller;

FIG. 6 illustrates a block diagram of an embodiment method; and

FIG. 7 illustrates an embodiment computer system.

Corresponding numerals and symbols in different figures generally referto corresponding parts unless otherwise indicated. The figures are drawnto clearly illustrate the relevant aspects of the preferred embodimentsand are not necessarily drawn to scale. To more clearly illustratecertain embodiments, a letter indicating variations of the samestructure, material, or process step may follow a figure number.

DETAILED DESCRIPTION OF ILLUSTRATIVE EMBODIMENTS

Illustrative embodiments of the system and method of the presentdisclosure are described below. In the interest of clarity, all featuresof an actual implementation may not be described in this specification.It will of course be appreciated that in the development of any suchactual embodiment, numerous implementation-specific decisions may bemade to achieve the developer's specific goals, such as compliance withsystem-related and business-related constraints, which will vary fromone implementation to another. Moreover, it should be appreciated thatsuch a development effort might be complex and time-consuming but wouldnevertheless be a routine undertaking for those of ordinary skill in theart having the benefit of this disclosure.

Reference may be made herein to the spatial relationships betweenvarious components and to the spatial orientation of various aspects ofcomponents as the devices are depicted in the attached drawings.However, as will be recognized by those skilled in the art after acomplete reading of the present disclosure, the devices, members,apparatuses, etc. described herein may be positioned in any desiredorientation. Thus, the use of terms such as “above,” “below,” “upper,”“lower,” or other like terms to describe a spatial relationship betweenvarious components or to describe the spatial orientation of aspects ofsuch components should be understood to describe a relative relationshipbetween the components or a spatial orientation of aspects of suchcomponents, respectively, as the device described herein may be orientedin any desired direction.

The increasing use of rotorcraft, in particular, for commercial andindustrial applications, has led to the development of larger morecomplex rotorcraft. However, as rotorcraft become larger and morecomplex, the differences between flying rotorcraft and fixed wingaircraft has become more pronounced. Since rotorcraft use one or moremain rotors to simultaneously provide lift, control attitude, controlaltitude, and provide lateral or positional movement, different flightparameters and controls are tightly coupled to each other, as theaerodynamic characteristics of the main rotors affect each control andmovement axis. For example, the flight characteristics of a rotorcraftat cruising speed or high speed may be significantly different than theflight characteristics at hover or at relatively low speeds.Additionally, different flight control inputs for different axes on themain rotor, such as cyclic inputs or collective inputs, affect otherflight controls or flight characteristics of the rotorcraft. Forexample, pitching the nose of a rotorcraft forward to increase forwardspeed will generally cause the rotorcraft to lose altitude. In such asituation, the collective may be increased to maintain level flight, butthe increase in collective causes increased power to the main rotorwhich, in turn, requires additional anti-torque force from the tailrotor. This is in contrast to fixed wing systems where the controlinputs are less closely tied to each other and flight characteristics indifferent speed regimes are more closely related to each other.

Recently, fly-by-wire (FBW) systems have been introduced in rotorcraftto assist pilots in stably flying the rotorcraft and to reduce workloadon the pilots. The FBW system may provide different controlcharacteristics or responses for cyclic, pedal or collective controlinput in the different flight regimes, and may provide stabilityassistance or enhancement by decoupling physical flight characteristicsso that a pilot is relieved from needing to compensate for some flightcommands issued to the rotorcraft. FBW systems may be implemented in oneor more flight control computers (FCCs) disposed between the pilotcontrols and flight control systems, providing corrections to flightcontrols that assist in operating the rotorcraft more efficiently orthat put the rotorcraft into a stable flight mode while still allowingthe pilot to override the FBW control inputs. The FBW systems in arotorcraft may, for example, automatically adjust power output by theengine to match a collective control input, apply collective or powercorrection during a cyclic control input, provide automation of one ormore flight control procedures, provide for default or suggested controlpositioning, or the like.

FBW systems for rotorcraft must provide stable flight characteristicsfor FBW controlled flight parameters while permitting the pilot tooverride or adjust any suggested flight parameters suggested by the FBWsystem. Additionally, in providing enhanced control and automatedfunctionality for rotorcraft flight, the FBW system must maintain anintuitive and easy to use flight control system for the pilot. Thus, theFBW system adjusts the pilot flight controls so that the controls are ina position associated with the relevant flight parameter. For example,the FBW system may adjust the collective stick to provide suggested orFBW controlled flight parameters, and which reflect a collective orpower setting. Thus, when the pilot releases the collective stick andthe FBW system provides collective control commands, the collectivestick is positioned intuitively in relation to the actual power orcollective setting so that, when the pilot grasps the collective stickto retake control, the control stick is positioned where the pilotexpects the stick to be positioned for the actual collective setting ofthe main rotor. Similarly, the FBW system use the cyclic stick to, forexample, adjust for turbulence, drift or other disturbance to the flightpath, and may move the cyclic stick as the FBW system compensates thecyclic control. Thus, when the pilot grasps the cyclic stick to takecontrol of flight from the FBW system, the cyclic stick is positioned toreflect the actual cyclic settings.

Embodiments of the present disclosure will be described with respect topreferred embodiments in a specific context, namely a system and methodfor smoothing a transition between a translational rate command (TRC)mode to a rate mode in a rotorcraft. Embodiments of the presentdisclosure may also be applied to other control mode transitions in theoperation and control of a rotorcraft.

In an embodiment of the present invention, a forward displacement on thecyclic controller or pilot stick of the rotorcraft is interpreted in oneof two ways. When the speed of the rotorcraft is 10 kts or less and thespeed command by displacement on the cyclic controller is less than 10kts, the displacement on the cyclic controller is interpreted as a speedcommand. For example, ½ inch of stick displacement may correspond to aspeed of 2 kts; 1″ of stick displacement may correspond to a speed of 4kts, and so on. However, when the speed of the rotorcraft is greaterthan 10 kts or the speed command by displacement on the cycliccontroller is greater than 10 kts, displacement on the cyclic controlleris interpreted as a rate command. For example, ½″ of stick displacementmay correspond to a rate of 10° per second; 1″ of stick displacement maycorrespond to a rate of 20° per second, and so on. By automaticallyswitching between the two different modes of operation based on thespeed of the rotorcraft, pilot workload may be reduced.

In some circumstances, however, the rotorcraft may undergo a physicaltransient during the transition between the speed control mode at lowspeeds and the rate control mode at higher speeds. This physicaltransient may be caused by sudden changes in flight control algorithmsand abrupt shifts in control signals. In embodiments of the presentinvention, the transition between speed control and rate control is madesmoother by controlling the gains and the signal path structure of anattitude controller used during the speed control mode. In one example,the overall gain of the attitude controller is reduced using a fader.During this time, a proportional path in the controller is maintainedwhile a value of an integrator is decremented to zero. Also during thistime the gain of a control path between the cyclic controller and therate control block is increased using a fader. By maintaining aproportional path in the attitude controller, decrementing theintegrator to zero, fading out the overall gain of the attitudecontroller and fading in the gain of the control path between the cycliccontroller and the rate control block, a smooth transition between speedcontrol mode and rate control mode may be made more smooth. It should beunderstood that embodiments of the present invention may be applied toother control signals and control paths of the rotorcraft.

FIG. 1 illustrates a rotorcraft 101 according to some embodiments. Therotorcraft 101 has a main rotor system 103, which includes a pluralityof main rotor blades 105. The pitch of each main rotor blade 105 may becontrolled by a swashplate 107 in order to selectively control theattitude, altitude and movement of the rotorcraft 101. The swashplate107 may be used to collectively and/or cyclically change the pitch ofthe main rotor blades 105. The rotorcraft 101 also has an anti-torquesystem, which may include a tail rotor 109, no-tail-rotor (NOTAR), ordual main rotor system. In rotorcraft with a tail rotor 109, the pitchof each tail rotor blade 111 is collectively changed in order to varythrust of the anti-torque system, providing directional control of therotorcraft 101. The pitch of the tail rotor blades 111 is changed by oneor more tail rotor actuators. In some embodiments, the FBW system sendselectrical signals to the tail rotor actuators or main rotor actuatorsto control flight of the rotorcraft.

Power is supplied to the main rotor system 103 and the anti-torquesystem by engines 115. There may be one or more engines 115, which maybe controlled according to signals from the FBW system. The output ofthe engine 115 is provided to a driveshaft 117, which is mechanicallyand operatively coupled to the rotor system 103 and the anti-torquesystem through a main rotor transmission 119 and a tail rotortransmission, respectively.

The rotorcraft 101 further includes a fuselage 125 and tail section 123.The tail section 123 may have other flight control devices such ashorizontal or vertical stabilizers, rudder, elevators, or other controlor stabilizing surfaces that are used to control or stabilize flight ofthe rotorcraft 101. The fuselage 125 includes a cockpit 127, whichincludes displays, controls, and instruments. It should be appreciatedthat even though rotorcraft 101 is depicted as having certainillustrated features, the rotorcraft 101 may have a variety ofimplementation-specific configurations. For instance, in someembodiments, cockpit 127 is configured to accommodate a pilot or a pilotand co-pilot, as illustrated. It is also contemplated, however, thatrotorcraft 101 may be operated remotely, in which case cockpit 127 couldbe configured as a fully functioning cockpit to accommodate a pilot (andpossibly a co-pilot as well) to provide for greater flexibility of use,or could be configured with a cockpit having limited functionality(e.g., a cockpit with accommodations for only one person who wouldfunction as the pilot operating perhaps with a remote co-pilot or whowould function as a co-pilot or back-up pilot with the primary pilotingfunctions being performed remotely). In yet other contemplatedembodiments, rotorcraft 101 could be configured as an unmanned vehicle,in which case cockpit 127 could be eliminated entirely in order to savespace and cost.

FIG. 2 illustrates a fly-by-wire flight control system 201 for arotorcraft according to some embodiments. A pilot may manipulate one ormore pilot flight controls in order to control flight of the rotorcraft.The pilot flight controls may include manual controls such as a cyclicstick 231 in a cyclic control assembly 217, a collective stick 233 in acollective control assembly 219, and pedals 239 in a pedal assembly 221.Inputs provided by the pilot to the pilot flight controls may betransmitted mechanically and/or electronically (e.g., via the FBW flightcontrol system) to flight control devices by the flight control system201. Flight control devices may represent devices operable to change theflight characteristics of the rotorcraft. Flight control devices on therotorcraft may include mechanical and/or electrical systems operable tochange the positions or angle of attack of the main rotor blades 105 andthe tail rotor blades in or to change the power output of the engines115, as examples. Flight control devices include systems such as theswashplate 107, tail rotor actuator 113, and systems operable to controlthe engines 115. The flight control system 201 may adjust the flightcontrol devices independently of the flight crew in order to stabilizethe rotorcraft, reduce workload of the flight crew, and the like. Theflight control system 201 includes engine control computers (ECCUs) 203,flight control computers 205, and aircraft sensors 207, whichcollectively adjust the flight control devices.

The flight control system 201 has one or more flight control computers205 (FCCs). In some embodiments, multiple FCCs 205 are provided forredundancy. One or more modules within the FCCs 205 may be partially orwholly embodied as software and/or hardware for performing anyfunctionality described herein. In embodiments where the flight controlsystem 201 is a FBW flight control system, the FCCs 205 may analyzepilot inputs and dispatch corresponding commands to the ECCUs 203, thetail rotor actuator 113, and/or actuators for the swashplate 107.Further, the FCCs 205 are configured and receive input commands from thepilot controls through sensors associated with each of the pilot flightcontrols. The input commands are received by measuring the positions ofthe pilot controls. The FCCs 205 also control tactile cueing commands tothe pilot controls or display information in instruments on, forexample, an instrument panel 241.

The ECCUs 203 control the engines 115. For example, the ECCUs 203 mayvary the output power of the engines 115 to control the rotational speedof the main rotor blades or the tail rotor blades. The ECCUs 203 maycontrol the output power of the engines 115 according to commands fromthe FCCs 205, or may do so based on feedback such a measured revolutionsper minute (RPM) of the main rotor blades.

The aircraft sensors 207 are in communication with the FCCs 205. Theaircraft sensors 207 may include sensors for measuring a variety ofrotorcraft systems, flight parameters, environmental conditions and thelike. For example, the aircraft sensors 207 may include sensors formeasuring airspeed, altitude, attitude, position, orientation,temperature, airspeed, vertical speed, and the like. Other sensors 207could include sensors relying upon data or signals originating externalto the rotorcraft, such as a global positioning system (GPS) sensor, aVHF Omnidirectional Range sensor, Instrument Landing System (ILS), andthe like.

The cyclic control assembly 217 is connected to a cyclic trim assembly229 having one or more cyclic position sensors 211, one or more cyclicdetent sensors 235, and one or more cyclic actuators or cyclic trimmotors 209. The cyclic position sensors 211 measure the position of thecyclic control stick 231. In some embodiments, the cyclic control stick231 is a single control stick that moves along two axes and permits apilot to control pitch, which is the vertical angle of the nose of therotorcraft and roll, which is the side-to-side angle of the rotorcraft.In some embodiments, the cyclic control assembly 217 has separate cyclicposition sensors 211 that measuring roll and pitch separately. Thecyclic position sensors 211 for detecting roll and pitch generate rolland pitch signals, respectively, (sometimes referred to as cycliclongitude and cyclic latitude signals, respectively) which are sent tothe FCCs 205, which controls the swashplate 107, engines 115, tail rotor109 or related flight control devices.

The cyclic trim motors 209 are connected to the FCCs 205, and receivesignals from the FCCs 205 to move the cyclic control stick 231. In someembodiments, the FCCs 205 determine a suggested cyclic stick positionfor the cyclic stick 231 according to one or more of the collectivestick position, the pedal position, the speed, altitude and attitude ofthe rotorcraft, the engine revolutions per minute (RPM), enginetemperature, main rotor RPM, engine torque or other rotorcraft systemconditions or flight conditions. The suggested cyclic stick position isa positon determined by the FCCs 205 to give a desired cyclic action. Insome embodiments, the FCCs 205 send a suggested cyclic stick positionsignal indicating the suggested cyclic stick position to the cyclic trimmotors 209. While the FCCs 205 may command the cyclic trim motors 209 tomove the cyclic stick 231 to a particular position (which would in turndrive actuators associated with swashplate 107 accordingly), the cyclicposition sensors 211 detect the actual position of the cyclic stick 231that is set by the cyclic trim motors 206 or input by the pilot,allowing the pilot to override the suggested cyclic stick position. Thecyclic trim motor 209 is connected to the cyclic stick 231 so that thepilot may move the cyclic stick 231 while the trim motor is driving thecyclic stick 231 to override the suggested cyclic stick position. Thus,in some embodiments, the FCCs 205 receive a signal from the cyclicposition sensors 211 indicating the actual cyclic stick position, and donot rely on the suggested cyclic stick position to command theswashplate 107.

Similar to the cyclic control assembly 217, the collective controlassembly 219 is connected to a collective trim assembly 225 having oneor more collective position sensors 215, one or more collective detentsensors 237, and one or more collective actuators or collective trimmotors 213. The collective position sensors 215 measure the position ofa collective control stick 233 in the collective control assembly 219.In some embodiments, the collective control stick 233 is a singlecontrol stick that moves along a single axis or with a lever typeaction. A collective position sensor 215 detects the position of thecollective control stick 233 and sends a collective position signal tothe FCCs 205, which controls engines 115, swashplate actuators, orrelated flight control devices according to the collective positionsignal to control the vertical movement of the rotorcraft. In someembodiments, the FCCs 205 may send a power command signal to the ECCUs203 and a collective command signal to the main rotor or swashplateactuators so that the angle of attack of the main blades is raised orlowered collectively, and the engine power is set to provide the neededpower to keep the main rotor RPM substantially constant.

The collective trim motor 213 is connected to the FCCs 205, and receivessignals from the FCCs 205 to move the collective control stick 233.Similar to the determination of the suggested cyclic stick position, insome embodiments, the FCCs 205 determine a suggested collective stickposition for the collective control stick 233 according to one or moreof the cyclic stick position, the pedal position, the speed, altitudeand attitude of the rotorcraft, the engine RPM, engine temperature, mainrotor RPM, engine torque or other rotorcraft system conditions or flightconditions. The FCCs 205 generate the suggested collective stickposition and send a corresponding suggested collective stick signal tothe collective trim motors 213 to move the collective stick 233 to aparticular position. The collective position sensors 215 detect theactual position of the collective stick 233 that is set by thecollective trim motor 213 or input by the pilot, allowing the pilot tooverride the suggested collective stick position.

The pedal control assembly 221 has one or more pedal sensors 227 thatmeasure the position of pedals or other input elements in the pedalcontrol assembly 221. In some embodiments, the pedal control assembly221 is free of a trim motor or actuator, and may have a mechanicalreturn element that centers the pedals when the pilot releases thepedals. In other embodiments, the pedal control assembly 221 has one ormore trim motors that drive the pedal to a suggested pedal positionaccording to a signal from the FCCs 205. The pedal sensor 227 detectsthe position of the pedals 239 and sends a pedal position signal to theFCCs 205, which controls the tail rotor 109 to cause the rotorcraft toyaw or rotate around a vertical axis.

The cyclic and collective trim motors 209 and 213 may drive the cyclicstick 231 and collective stick 233, respectively, to suggestedpositions. The cyclic and collective trim motors 209 and 213 may drivethe cyclic stick 231 and collective stick 233, respectively, tosuggested positions, but this movement capability may also be used toprovide tactile cueing to a pilot. The trim motors 209 and 213 may pushthe respective stick in a particular direction when the pilot is movingthe stick to indicate a particular condition. Since the FBW systemmechanically disconnects the stick from one or more flight controldevices, a pilot may not feel a hard stop, vibration, or other tactilecue that would be inherent in a stick that is mechanically connected toa flight control assembly. In some embodiments, the FCCs 205 may causethe trim motors 209 and 213 to push against a pilot command so that thepilot feels a resistive force, or may command one or more frictiondevices to provide friction that is felt when the pilot moves the stick.Thus, the FCCs 205 control the feel of a stick by providing pressureand/or friction on the stick.

Additionally, the cyclic control assembly 217, collective controlassembly 219 and/or pedal control assembly 221 may each have one or moredetent sensors that determine whether the pilot is handling a particularcontrol device. For example, the cyclic control assembly 217 may have acyclic detent sensor 235 that determines that the pilot is eitherholding or moving the cyclic stick 231, while the collective controlassembly 219 has a collective detent sensor 237 that determines whetherthe pilot is either holding or moving the collective stick 233. Thesedetent sensors 235, 237 detect motion and/or position of the respectivecontrol stick that is caused by pilot input, as opposed to motion and/orposition caused by commands from the FCCs 205, rotorcraft vibration, andthe like and provide feedback signals indicative of such to the FCCs.When the FCCs 205 detect that a pilot has control of, or ismanipulating, a particular control, the FCCs 205 may determine thatstick to be out-of-detent (00D). Likewise, the FCCs may determine thatthe stick is in-detent (ID) when the signals from the detent sensorsindicate to the FCCs 205 that the pilot has released a particular stick.The FCCs 205 may provide different default control or automated commandsto one or more flight systems based on the detent status of a particularstick or pilot control.

Moving now to the operational aspects of flight control system 201, FIG.3 illustrates in a highly schematic fashion, a manner in which flightcontrol system 201 may implement FBW functions as a series ofinter-related feedback loops running certain control laws. FIG. 3representatively illustrates a three-loop flight control system 201according to an embodiment. In some embodiments, elements of thethree-loop flight control system 201 may be implemented at leastpartially by FCCs 205. As shown in FIG. 3, however, all, some, or noneof the components (301, 303, 305, 307) of three-loop flight controlsystem 201 could be located external or remote from the rotorcraft 100and communicate to on-board devices through a network connection 309.

The three-loop flight control system 201 of FIG. 3 has a pilot input311, an outer loop 313, a rate (middle) loop 315, an inner loop 317, adecoupler 319, and aircraft equipment 321 (corresponding, e.g., toflight control devices such as swashplate 107, tail rotor transmission212, etc., to actuators (not shown) driving the flight control devices,to sensors such as aircraft sensors 207, position sensors 211, 215,detent sensors 235, 237, etc., and the like).

In the example of FIG. 3, a three-loop design separates the innerstabilization and rate feedback loops from outer guidance and trackingloops. The control law structure primarily assigns the overallstabilization task and related tasks of reducing pilot workload to innerloop 317. Next, middle loop 315 provides rate augmentation. Outer loop313 focuses on guidance and tracking tasks. Since inner loop 317 andrate loop 315 provide most of the stabilization, less control effort isrequired at the outer loop level. As representatively illustrated inFIG. 3, a switch 322 may be provided to turn outer loop flightaugmentation on and off, as the tasks of outer loop 313 are notnecessary for flight stabilization.

In some embodiments, the inner loop 317 and rate loop 315 include a setof gains and filters applied to roll/pitch/yaw 3-axis rate gyro andacceleration feedback sensors. Both the inner loop and rate loop maystay active, independent of various outer loop hold modes. Outer loop313 may include cascaded layers of loops, including an attitude loop, aspeed loop, a position loop, a vertical speed loop, an altitude loop,and a heading loop. In accordance with some embodiments, the controllaws running in the illustrated the loops allow for decoupling ofotherwise coupled flight characteristics, which in turn may provide formore stable flight characteristics and reduced pilot workload.Furthermore, the outer loop 313 may allow for automated orsemi-automated operation of certain high-level tasks or flight patterns,thus further relieving the pilot workload and allowing the pilot tofocus on other matters including observation of the surrounding terrain.

FIG. 4 illustrates a flight control system 400 according to anembodiment of the present invention. Pilot stick pitch block 401 a andpilot stick roll block 401 b represent, for example, the respectivepitch and roll commands emanating from cyclic controller 217 of therotorcraft. As shown, pilot stick blocks 401 a and 401 b interface toflight controller 402. In various embodiments, flight controller 402 isimplemented using a flight computer, or other processing hardware.Flight controller 402 also interfaces with and controls aircraftequipment 321 representing various actuators, sensors, and the physicalbody of the rotorcraft. In various embodiments, flight controller 402controls aircraft equipment 321 using three loops: an inner loop, a ratefeedback loop and a state feedback loop: the inner loop stabilizes thedynamics of the rotorcraft, the rate loop controls the angular rates ofthe rotor craft, and the outer loop provides control signals to theinner loop and/or rate loops to effect a desired attitude, speed andposition of the rotorcraft. In some embodiments, the outer loop supportsand provides flight augmentation or auto-pilot functionality and may bemanually or automatically disabled based on flight and systemconditions. The inner loop and rate feedback loops, on the other hand,remain operational to provide stability to the rotorcraft.

For purposes of illustration, flight controller 402 is illustrated withrespect to the general control blocks that affect the rotational rate ofan embodiment rotorcraft, namely the control blocks affecting the rollrate and the pitch rate of the rotorcraft. It should be understood thatflight controller 402 may also include other controllers and controlpaths that affect the yaw and other states of the rotorcraft in additionto the pitch rate and roll rate. As shown, the inner stabilization loopis controlled by inner loop controller 317, the rate loop is controlledby pitch rate controller 315 a and roll rate controller 315 b, and theouter loop is controlled by outer loop controller 313 in conjunctionwith TRC mode attitude controllers 408 a and 408 b that provide TRC modepitch and roll control, respectively, during embodiment TRC modes.

Each of inner loop controller 317, decoupler 319 and rate controller 315may be implemented using flight control algorithms known in the art.Inner loop controller 317 receives sensor feedback from sensors such asgyroscopes and accelerometers within the rotorcraft and provides controlsignals to various actuators, such as swashplate 107 to stabilize therotorcraft. Rate controller 315 receives rate feedback from rategyroscopes on all axes and provides a rate command signal to inner loopcontroller 317 based on the rate feedback and the position of pilotstick pitch block 401 a and pilot stick roll block 401 b in some modesof operation. Decoupler 319 receives the various rate commands andapproximately decouples all 4-axes (pitch, roll, yaw, and vertical) suchthat, for example, a forward longitudinal stick input does not requirethe pilot to push the stick diagonally.

Outer loop controller 313 receives state feedback from the sensors ofaircraft equipment 321. This state feedback may include, for example,speed, position and attitude. In a translational rate commend (TRC)mode, outer loop controller 313 receives a command from the pilot stickrepresented by pilot stick pitch block 401 a and pilot stick roll block401 b, determines a corresponding translational speed based on the pilotstick command, and determines corresponding pitch and roll attitudecommands based on the determined translational speed. The translationalspeed may include a forward component based on commands from pilot stickpitch block 401 a and a lateral component based on commands from pilotstick roll block 401 b. In some embodiments, the correspondingtranslational speed is determined by multiplying the pilot stickcommand, or a vector sum of the pitch and roll pilot stick commands witha scale factor and/or by using a lookup table. The translational speedmay be calculated using a vector sum as follows:

V _(T) =k√{square root over (S _(LON) ² +S _(LAT) ²)},

where VT is the translational speed, k is a scale factor, S_(LON) is thepilot stick pitch command, and S_(LAT) is the pilot stick roll command.In some embodiments, the vector sum is not used and the translationalspeed is based on the forward speed and/or the lateral speed.

The pitch and roll attitude commands may be determined from thetranslational speed using a speed control loop. During the TRC mode, TRCmode attitude controllers 408 a and 408 b individually calculate pitchand roll attitude errors by subtracting the pitch attitude feedback fromthe pitch attitude command and by subtracting the roll attitude feedbackfrom the roll attitude command. In various embodiments, the pitch androll attitude feedback is a component of the state feedback. TRC modeattitude controller 408 a applies a dynamic control algorithm to thepitch attitude error to produce an outer loop pitch rate command, andTRC mode attitude controller 408 b applies a dynamic control algorithmto the roll attitude error to produce an outer loop roll rate command.These outer loop rate commands are applied to decoupler 319 via summingblocks 412 a and 412 b. While the embodiment of FIG. 4 shows TRC modeattitude controllers 408 a and 408 b used to provide TRC mode control toboth the pitch and roll channels, it should be understood that inalternative embodiments, TRC mode control may be applied to a singleattitude channel, such as the pitch channel for speed control in theforward direction.

In various embodiments of the present invention, the TRC mode isautomatically selected based on the velocity or ground speed of therotorcraft. This ground speed may be measured, for example, using GPS.In one example, the TRC mode is selected when the ground speed is lessthan 10 kts, in which a fixed offset of the pilot stick effects aconstant translation. When the ground speed is 10 kts and greater, thepilot stick effects a rate in a rate control mode. In the embodiment ofFIG. 4, the TRC mode is activated by enabling TRC mode attitudecontroller 408 a and 408 b when the ground speed is less than 10 kts anddecoupling the pilot stick pitch block 401 a from pitch rate controller315 a via fader 404 a and decoupling the pilot stick roll block 401 bfrom the roll rate controller 315 b via fader 404 b. It should beappreciated that the 10 kts threshold is just one example of manypossible groundspeed thresholds. In alternative embodiments, otherthreshold values may be used. Likewise, the rate control mode isselected by coupling pilot stick pitch block 401 a to pitch ratecontroller 315 a via fader 404 a, and by coupling pilot stick roll block401 b to roll rate controller 315 a to pitch rate controller via fader404 b. In some embodiments, pilot stick blocks 401 a and 401 b arecoupled to rate controllers 315 a and 315 b when the pilot stick is outof detent. In some embodiments, hysteresis may be applied to the groundspeed threshold to prevent metastability of the control mode when therotorcraft is flying at about 10 kts. In such an embodiment the groundspeed threshold for transitioning from the TRC to the rate control modeis greater than the ground speed threshold for transitioning from therate control mode to the TRC mode. For example, in one embodiment, TRCmode attitude controllers 408 a and 408 b are disabled when the pilotstick command is greater than 10 kts and then re-enabled when the pilotstick command is less than 3 kts. Alternatively, otherhysteresis/threshold values may be used. Alternatively, a singlethreshold may be used without hysteresis. In various embodiments, the 10kts and 3 kts speed thresholds may be applied to the forward speed, thelateral speed or a combination of the forward and lateral speed. Such acombination may be determined, for example, by calculating a vector sumof the forward speed and the lateral speed.

During operation, faders 404 a and 404 b provide a gain of one when thepilot stick is out of detent, and provides a gain of zero when the pilotstick is in detent. In various embodiments, the gain of faders 404 a and404 b linearly increase or decrease between zero and one over apredetermined period of time. This predetermined period of time isbetween about 1 s and about 10 s; however, fading times outside of thisrange may be implemented depending on the particular embodiment and itsspecifications. In further alternative embodiments, the gains of fader404 a and 404 b may be different from one and zero. In some embodiments,pitch rate controllers 315 a and roll rate controller 315 b each ahigh-pass filter and/or washout filter coupled to their respectiveinputs. Thus, when the stick is out of detent and held at a steadydisplacement, the effective command produced by rate controller 315 aand 315 b as a result of a pilot stick command goes to zero. Thishigh-pass filter or washout filter also prevents the output of TRC modeattitude controller 408 a and 408 b and the output of faders 404 a and404 b from conflicting or “fighting” with each other when the pilotstick is out of detent.

FIG. 5 illustrates a block diagram of TRC mode attitude controller 408according to an embodiment of the present invention that may be used toimplement TRC attitude controllers 408 a and 408 b illustrated in Figure4. As shown, TRC mode attitude controller 408 includes subtractor 430,fader 432, dynamic compensation block 434 and proportional-integral (PI)controller 450. Subtractor 430 produces an error signal based on theattitude command produced by the outer loop controller 313 (FIG. 4) anddynamic compensation block 434 compensates for the dynamics of therotorcraft in order to improve stability and/or adjust the time responseof the attitude loop. Dynamic compensation block 434 may include variouscontrol blocks known in the art including but not limited a PIDcontroller and a lead-lag compensator.

PI controller 450 includes a proportional signal path that includesproportional gain block 436 coupled in parallel with an integral signalpath that includes integral gain block 438 and integrator 440. Invarious embodiments, proportional gain block 436 and integral gain block438 may be implemented, for example, by performing a multiplication orscaling operation. Integrator 440 may be implemented, for example, usingan accumulator.

During TRC mode operation below the 10 kts speed threshold, fader 432has a gain of one and the input of integrator 440 is coupled to theoutput of dynamic compensation block 434 via integral gain block 438. Onthe other hand, in the rate control mode in which the rotorcraft has aground speed of 10 kts and above or the speed command by displacement onthe cyclic controller is greater than 10 kts, fader 432 has a gain ofzero, and the input of integrator 440 is coupled to its output viafeedback gain block 444 and limiter 442, which effectively causesintegrator 440 to decrement to zero or to a DC value representing zerooutput. The combination of zero fader gain and zero integrator outputeffectively disables TRC mode attitude controller 408. When the groundspeed of the rotorcraft exceeds the 10 kts threshold, the gain of fader432 linearly decreases to zero over a predetermined time period and theinput of the integrator is switched from the output of integral gainblock 438 to the output of limiter 442. Once the input to integrator isswitched to the output of limiter 442 via switch 452, the feedbackaction of the loop formed by integrator 440, feedback gain 444 andlimiter 442 forces the output of integrator 440 decrement to zero over aperiod of time. Limiter 442 limits the rate at which integrator 440 isdecremented such that the decay of integrator 440 has more of a rampresponse instead of an exponential response for high integrator outputvalues. Thus, in some embodiments, during a mode transition from TRCmode to rate control mode, the actuator command generated by summer 446produces a gently decreasing or decaying signal from the forward path ofthe TRC mode attitude controller 408 due to fader 432 that is summedwith a gently decreasing or decaying integrator value produced byintegrator 440. At the same time as the output of the TRC mode attitudecontroller 408 is decaying, the pilot stick command to rate controller315 is increasing due the increasing gain of fader 404. The net resultof the decreasing output of the TRC mode attitude controller 408 and theincreasing output of fader 404 a or 404 b to rate controller 315 is asmooth transition from TRC mode to attitude mode in some embodiments. Invarious embodiments, the time that it takes for faders 404 a, 404 b and432 to change their gains may be the same or may be different from eachother. In some embodiments, the fade-in and fade-out times of faders 404a, 404 b and 432 may be the same or they may be different from eachother.

Similarly, during a mode transition from rate mode to TRC mode, thepilot stick control signal produced by fader 404 a or 404 b decreaseswhile the forward path of TRC mode attitude controller 408 increases dueto the gain of fader 432 increasing and due to the integral path of PIcontroller 450 being activated by coupling the input of integrator 440to the output of integral gain block 438 via switch 452.

In alternative embodiments of the present invention, integrator may bedecremented in a different manner. For example, instead of usingfeedback gain 444 to decrease the output of the integrator, a fader maybe coupled to the output of integrator 440 or may be coupled aftersummer 446.

In various embodiments, the transfer function and implementation ofdynamic compensation block 434, the gains of fader 432, proportionalgain 435, integral gain 438 and feedback gain 444, and the static anddynamic behavior of fader 432, limiter 442 and integrator 440 may beadjusted according to the specification and requirements of therotorcraft and may have different implementations and values with regardto the roll, pitch and other attitude channels.

It should be understood that the transition between TRC and rate controlmode with respect to the various rate commands is just one example ofmany possible system configurations that utilize embodiment modeswitching systems and methods. In alternative embodiments, other controlchannels besides the pitch rate and roll rate control channel may becontrolled. For example, embodiments of the present invention may beused to transition between attitude hold/command mode and a speedhold/command mode, transition between attitude hold/command mode and aposition hold/command mode, transition between direct verticalcollective control mode and an altitude hold/command mode, and/ortransition between vertical speed hold/command mode and an altitudehold/command mode. It should be appreciated that these are only a fewexamples of many possible applications.

FIG. 6 illustrates a block diagram of an embodiment method 500 ofoperating a rotorcraft that may be executed, for example, by a flightcomputer of a fly-by-wire system. In an embodiment, the pilot stickcommand of the cyclic control is interpreted as a translational ratecommand in step 518 in which the flight computer controls the rotorcraftto have a translational velocity that is proportional to a physicaloffset of the collective control. During step 518, a dynamic controlleradjusts the attitude of the rotorcraft in order to maintain the desiredtranslational velocity. When the velocity of the rotorcraft exceeds athreshold V_(threshold2) according to step 520 or when the pilot stickcommand exceeds V_(threshold2), the operational mode of the flightcomputer transitions to a rate command mode in step 508 in which thepilot stick command of the cyclic control is interpreted as a ratecommand. In this mode, the flight computer controls the rotorcraft tohave a rate that is proportional to the physical offset of thecollective control.

In order to provide a smooth transition from the translational commandmode of step 518 to the rate command mode of step 508 and to reduce theoccurrence of physical transients of the rotorcraft, steps 502, 504 and506 are executed by the flight computer. In step 502, the gain of afirst dynamic controller, which may be implemented, for example, asouter loop controller that provides a rate control in response to adesired attitude command, is faded out over first period of time. Instep 504, a state of an integrator of the first dynamic controller isdecreases over a second period of time, and in step 506 a gain of thepilot stick command is faded in over a third period of time. In someembodiments, steps 502, 504 and 506 occur concurrently such that thegain of the first dynamic controller is reduced at the same time as thegain of the pilot stick command to the rate controller is faded in.During this time, the value of the integrator is reduced. In someembodiments, the value of the integrator is decremented by disconnectingthe integrator from the input of the first dynamic controller andcoupling the input of the integrator to its output. The input and/oroutput value of the integrator may be limited while the integrator valueis reduced in order to increase the amount of time that it takes todecrement the value of the integrator to zero.

The rate command continues to be determined by the pilot stick commandin step 508 until the velocity of the rotorcraft is less than or equalto threshold V_(threshold1) and the stick command S_(command) is lessthan an offset S_(threshold) that represents a particular translationalvelocity command according to step 510. In some embodiments,S_(threshold) may be set to a pilot stick offset that is less than 3kts. Alternatively, other thresholds may be used. Once this condition isdetected, the operational mode of the flight computer transitions to thetranslational rate command mode of step 518 via steps 512, 514 and 516.In some embodiments, threshold V_(threshold1) is equal to thresholdV_(threshold2), while in other embodiments, V_(threshold2) is greaterthreshold V_(threshold1) in order to provide hysteresis. In steps 512,the gain of the first dynamic controller is faded in, in step 514, theintegrator of the first dynamic controller is reactivated by couplingits input to the signal path of the first dynamic controller, and instep 516 the gain of the pilot stick command to the rate controller isfaded out. In some embodiments, steps 512, 514 and 516 occurconcurrently such that the gain of the first dynamic controller isincreased at the same time as the gain of the pilot stick command to therate controller is reduced and at the same time the integrator of thefirst dynamic controller is reactivated.

FIG. 7 illustrates a computer system 601. The computer system 601 can beconfigured for performing one or more functions with regard to theoperation of the flight control system 201 and the method 500, asdescribed herein. Further, any processing and analysis can be partly orfully performed by the computer system 601. The computer system 601 canbe partly or fully integrated with other aircraft computer systems orcan be partly or fully removed from the rotorcraft.

The computer system 601 can include an input/output (I/O) interface 603,an analysis engine 605, and a database 607. Alternative embodiments cancombine or distribute the I/O interface 603, the analysis engine 605,and the database 607, as desired. Embodiments of the computer system 601may include one or more computers that include one or more processorsand memories configured for performing tasks described herein. This caninclude, for example, a computer having a central processing unit (CPU)and non-volatile memory that stores software instructions forinstructing the CPU to perform at least some of the tasks describedherein. This can also include, for example, two or more computers thatare in communication via a computer network, where one or more of thecomputers include a CPU and non-volatile memory, and one or more of thecomputer's non-volatile memory stores software instructions forinstructing any of the CPU(s) to perform any of the tasks describedherein. Thus, while the exemplary embodiment is described in terms of adiscrete machine, it should be appreciated that this description isnon-limiting, and that the present description applies equally tonumerous other arrangements involving one or more machines performingtasks distributed in any way among the one or more machines. It shouldalso be appreciated that such machines need not be dedicated toperforming tasks described herein, but instead can be multi-purposemachines, for example computer workstations, that are suitable for alsoperforming other tasks.

The I/O interface 603 can provide a communication link between externalusers, systems, and data sources and components of the computer system601. The I/O interface 603 can be configured for allowing one or moreusers to input information to the computer system 601 via any knowninput device. Examples can include a keyboard, mouse, touch screen,and/or any other desired input device. The I/O interface 603 can beconfigured for allowing one or more users to receive information outputfrom the computer system 601 via any known output device. Examples caninclude a display monitor, a printer, cockpit display, and/or any otherdesired output device. The I/O interface 603 can be configured forallowing other systems to communicate with the computer system 601. Forexample, the I/O interface 603 can allow one or more remote computer(s)to access information, input information, and/or remotely instruct thecomputer system 601 to perform one or more of the tasks describedherein. The I/O interface 603 can be configured for allowingcommunication with one or more remote data sources. For example, the I/Ointerface 603 can allow one or more remote data source(s) to accessinformation, input information, and/or remotely instruct the computersystem 601 to perform one or more of the tasks described herein.

The database 607 provides persistent data storage for the computersystem 601. Although the term “database” is primarily used, a memory orother suitable data storage arrangement may provide the functionality ofthe database 607. In alternative embodiments, the database 607 can beintegral to or separate from the computer system 601 and can operate onone or more computers. The database 607 preferably provides non-volatiledata storage for any information suitable to support the operation ofthe flight control system 201 and the method 500, including varioustypes of data discussed further herein. The analysis engine 605 caninclude various combinations of one or more processors, memories, andsoftware components.

Embodiments of the present invention are summarized here. Otherembodiments can also be understood from the entirety of thespecification and the claims filed herein. One general aspect includes amethod of operating a rotorcraft that includes operating the rotorcraftin a first mode including determining a translational speed based on apilot stick signal generated by a pilot stick assembly, determining anattitude based on the determined translational speed, determining anactuator command based on the determined attitude by using a firstdynamic controller having an integrator, and providing an output of thefirst dynamic controller to an actuator, where a translational speed ofthe rotorcraft is proportional to an amplitude of the pilot stick signalin the first mode. The method also includes transitioning from the firstmode to a second mode when a velocity of the rotorcraft exceeds a firstvelocity threshold, where transitioning includes fading out a gain ofthe first dynamic controller over a first period of time and decreasinga value of the integrator over a second period of time. The method alsoincludes operating the rotorcraft in the second mode including providingan output of an attitude rate controller to the actuator, where theoutput of the attitude rate controller is proportional to an amplitudeof the pilot stick signal.

Implementations may include one or more of the following features. Themethod where: determining the translational speed based on a pilot sticksignal includes determining a forward speed component based on cycliclongitude stick control portion of the pilot stick signal anddetermining a lateral speed component based on a cyclic latitude stickcontrol portion of the pilot stick signal; determining the attitudebased on the determined translational speed includes determining a pitchattitude based on the determined forward speed component and a rollattitude based on the determined lateral speed component; determiningthe actuator command includes determining a pitch actuator command basedon the determined pitch attitude and a roll actuator command based onthe determined roll attitude; the first dynamic controller includes afirst pitch dynamic controller having a pitch integrator, and a firstroll dynamic controller having a roll integrator; providing the outputof the first dynamic controller to the actuator includes providing anoutput of the first pitch dynamic controller to a pitch actuator and anoutput of the first roll dynamic controller to a roll actuator; fadingout the gain of the first dynamic controller over a first period of timeincludes fading out a gain of the first pitch dynamic controller and thefirst roll dynamic controller over the first period of time; anddecreasing the value of the integrator over the second period of timeincludes decreasing a value of the pitch integrator and the rollintegrator over the second period of time.

In an embodiment, the translational speed further includes determining avector sum of the forward speed component and the lateral speedcomponent. In some embodiments, the translational speed includes aforward speed, the attitude includes a pitch attitude, the attitude ratecontroller includes a pitch rate controller and the actuator includes apitch actuator. The method may further including transitioning from thesecond mode to the first mode when the velocity of the rotorcraftdecreases below a second velocity threshold, where transitioning fromthe second mode to the first mode includes fading in the gain of thefirst dynamic controller over a third period of time and reactivatingthe integrator. In some embodiments, the velocity may be a ground speed.

The method may further included fading in a gain of the pilot sticksignal in the attitude rate controller over a fourth period of time whena pilot stick of the pilot stick assembly is out of detent; and fadingout a gain of the pilot stick signal in the attitude rate controllerover a fifth period of time when the pilot stick of the pilot stickassembly is in detent. In some embodiments, the first period of time isa same length as the third period of time, the second period of time isa same length as the fifth period of time, and the first velocitythreshold is the same as the second velocity threshold. In someembodiments, the first period of time and the second period of time arebetween 1 s and 10 s.

In an embodiment, the first dynamic controller further includes aproportional path, and an output of the first dynamic controller is asum of the proportional path and the output of the integrator.Decreasing the value of the integrator may include coupling the outputof the integrator to an input of the integrator via a first feedbackpath. In some embodiments, the first feedback path includes a limiter.

Another general aspect includes a flight control system for a rotorcraftincluding: a processor and a non-transitory computer readable storagemedium with an executable program stored thereon, where the executableprogram includes instructions to: receive a pilot control signal via afirst interface of the processor; in a first mode determine a firstvalue based on the received pilot control signal, determine a secondvalue based on the determined first value, determine an actuator commandbased on the determined first value, where determining the actuatorcommand includes executing a first dynamic controller that has anintegrator, and providing an output of the first dynamic controller toan actuator via a second interface of the processor, where a state ofthe rotorcraft corresponding to the first value is configured to beproportional to the received pilot control signal. The method furtherincludes transitioning from the first mode to a second mode when a firstcondition of the rotorcraft crosses a first predetermined threshold,where transitioning includes fading out a gain of the first dynamiccontroller over a first period of time, and decreasing a value of theintegrator over a second period of time; and in the second mode,providing an output of a second dynamic controller to the actuator viathe second interface of the processor, where the output of the seconddynamic controller is proportional to the received pilot control signal.

Implementations may include one or more of the following features. Theflight control system where the first value is a translational speed,the second value is an attitude, and the first condition of therotorcraft is a velocity of the rotorcraft. In some embodiments, thefirst value is a translational speed, the second value is an attitude,and the first condition of the rotorcraft is a velocity of therotorcraft. The attitude may include a pitch attitude and a rollattitude, executing the first dynamic controller may include executing afirst pitch dynamic controller having a pitch integrator and executing afirst roll dynamic controller having a roll integrator. Providing theoutput of the first dynamic controller to the actuator via a secondinterface of the processor may include providing an output of the firstpitch dynamic controller to a pitch actuator and providing an output ofthe first roll dynamic controller to a roll actuator.

In an embodiment, the executable program further includes instructionsto fade in a gain of the pilot control signal in the second dynamiccontroller when a pilot control stick is out of detent; and fade out again of the pilot control signal in the second dynamic controller whenthe pilot control stick is in detent. The executable program may furtherinclude instructions to transition from the second mode to the firstmode when the first condition of the rotorcraft decreases below a secondpredetermined threshold, where transitioning from the second mode to thefirst mode includes fading in the gain of the first dynamic controllerover a third period of time, and reactivating the integrator. In someembodiments, the first predetermined threshold is greater than thesecond predetermined threshold.

In an embodiment, the first dynamic controller further includes aproportional path, and an output of the first dynamic controller is asum of the proportional path and the output of the integrator.Decreasing the value of the integrator may include coupling the outputof the integrator to an input of the integrator via a first feedbackpath. In some example, the first feedback path includes a limiter. Thefirst predetermined threshold may include a ground speed and the firstperiod of time and the second period of time may be between 1 s and 10s.

Another general aspect includes a rotorcraft including: a body; a powertrain coupled to the body and including a power source and a drive shaftcoupled to the power source; a rotor system coupled to the power trainand including a plurality of rotor blades; a flight control systemoperable to change at least one operating condition of the rotor system;a pilot control assembly configured to receive commands from a pilot,where the flight control system is a fly-by-wire flight control systemin electrical communication with the pilot control assembly; and aflight control computer in electrical communication between the flightcontrol system and the pilot control assembly. The flight controlcomputer configured to: receive, from the pilot control assembly a pilotcommand to change a first flight characteristic, when a velocity of therotorcraft is less than a first velocity threshold, interpret the firstflight characteristic as a translational speed in a first mode anddetermine an attitude based on the translational speed via using anattitude controller including an integrator and a proportional pathparallel to the integrator. When the velocity of the rotorcraft isgreater than a second velocity threshold, the first flightcharacteristic is interpreted as a rate in a second mode, and when thevelocity of the rotorcraft increases past the second velocity threshold,fade out a gain of the attitude controller is faded out, and a value ofthe integrator is successively decreased.

Implementations may include one or more of the following features. Therotorcraft where: the attitude includes a pitch attitude; and the rateincludes a pitch rate. In the attitude further includes a roll attitude;and the rate further includes a roll rate. The flight control computermay be further configured to transition from the second mode to thefirst mode when the velocity of the rotorcraft decreases below the firstvelocity threshold, and transitioning from the second mode to the firstmode may include fading in the gain of the attitude controller, andreactivating the integrator. In some embodiments, the first velocitythreshold is less than the second velocity threshold.

In an embodiment, the flight control computer is further configured to:fade in a gain of the pilot command to an attitude rate controller whena pilot stick of the pilot control assembly is out of detent; and fadeout a gain of the pilot command to the attitude rate controller when thepilot stick of the pilot control assembly is in detent.

Advantages of embodiments include the ability to automatically andsmoothly transition between a TRC mode and a rate control mode orbetween a translational speed hold mode and a rotational rate controlmode. Further advantages include making the transition smooth andseamless enough that little change in attitude is noticeable to pilot oroccupants of the aircraft.

While this invention has been described with reference to illustrativeembodiments, this description is not intended to be construed in alimiting sense. Various modifications and combinations of theillustrative embodiments, as well as other embodiments of the invention,will be apparent to persons skilled in the art upon reference to thedescription. It is therefore intended that the appended claims encompassany such modifications or embodiments.

What is claimed is:
 1. A method of operating a rotorcraft, the methodcomprising: operating the rotorcraft in a first mode comprisingdetermining a translational speed based on a pilot stick signalgenerated by a pilot stick assembly, determining an attitude based onthe determined translational speed, determining an actuator commandbased on the determined attitude, determining the actuator commandcomprising using a first dynamic controller having an integrator;providing an output of the first dynamic controller to an actuator,wherein a translational speed of the rotorcraft is proportional to anamplitude of the pilot stick signal in the first mode; and transitioningfrom the first mode to a second mode when a velocity of the rotorcraftexceeds a first velocity threshold, transitioning comprising fading outa gain of the first dynamic controller over a first period of time, anddecreasing a value of the integrator over a second period of time; andoperating the rotorcraft in the second mode comprising providing anoutput of an attitude rate controller to the actuator, wherein theoutput of the attitude rate controller is proportional to an amplitudeof the pilot stick signal.
 2. The method of claim 1, wherein:determining the translational speed based on a pilot stick signalcomprises determining a forward speed component based on cycliclongitude stick control portion of the pilot stick signal anddetermining a lateral speed component based on a cyclic latitude stickcontrol portion of the pilot stick signal; determining the attitudebased on the determined translational speed comprises determining apitch attitude based on the determined forward speed component and aroll attitude based on the determined lateral speed component;determining the actuator command comprises determining a pitch actuatorcommand based on the determined pitch attitude and a roll actuatorcommand based on the determined roll attitude; the first dynamiccontroller comprises a first pitch dynamic controller having a pitchintegrator, and a first roll dynamic controller having a rollintegrator; providing the output of the first dynamic controller to theactuator comprises providing an output of the first pitch dynamiccontroller to a pitch actuator and an output of the first roll dynamiccontroller to a roll actuator; fading out the gain of the first dynamiccontroller over a first period of time comprises fading out a gain ofthe first pitch dynamic controller and the first roll dynamic controllerover the first period of time; and decreasing the value of theintegrator over the second period of time comprises decreasing a valueof the pitch integrator and the roll integrator over the second periodof time.
 3. The method of claim 2, wherein determining the translationalspeed further comprises determining a vector sum of the forward speedcomponent and the lateral speed component.
 4. The method of claim 1,wherein: the translational speed comprises a forward speed; the attitudecomprises a pitch attitude; the attitude rate controller comprises apitch rate controller; and the actuator comprises a pitch actuator. 5.The method of claim 1, further comprising transitioning from the secondmode to the first mode when the velocity of the rotorcraft decreasesbelow a second velocity threshold, transitioning from the second mode tothe first mode comprising: fading in the gain of the first dynamiccontroller over a third period of time, and reactivating the integrator.6. The method of claim 5, further comprising: fading in a gain of thepilot stick signal in the attitude rate controller over a fourth periodof time when a pilot stick of the pilot stick assembly is out of detent;and fading out a gain of the pilot stick signal in the attitude ratecontroller over a fifth period of time when the pilot stick of the pilotstick assembly is in detent.
 7. The method of claim 6, wherein: thefirst period of time is a same length as the third period of time; thesecond period of time is a same length as the fifth period of time; andthe first velocity threshold is the same as the second velocitythreshold.
 8. The method of claim 5, wherein the first velocitythreshold is greater than the second velocity threshold.
 9. The methodof claim 1, wherein the first dynamic controller further comprises aproportional path, and an output of the first dynamic controller is asum of the proportional path and the output of the integrator.
 10. Themethod of claim 1, wherein decreasing the value of the integratorcomprises coupling the output of the integrator to an input of theintegrator via a first feedback path.
 11. The method of claim 10,wherein the first feedback path comprises a limiter.
 12. The method ofclaim 1, wherein the velocity of the rotorcraft comprises a groundspeed.
 13. The method of claim 1, wherein the first period of time andthe second period of time are between 1 s and 10 s.
 14. A flight controlsystem for a rotorcraft comprising: a processor and a non-transitorycomputer readable storage medium with an executable program storedthereon, the executable program including instructions to: receive apilot control signal via a first interface of the processor; in a firstmode determine a first value based on the received pilot control signal,determine a second value based on the determined first value, determinean actuator command based on the determined first value, determining theactuator command comprising executing a first dynamic controller thathas an integrator, and providing an output of the first dynamiccontroller to an actuator via a second interface of the processor,wherein a state of the rotorcraft corresponding to the first value isconfigured to be proportional to the received pilot control signal; andtransitioning from the first mode to a second mode when a firstcondition of the rotorcraft crosses a first predetermined threshold,transitioning comprising fading out a gain of the first dynamiccontroller over a first period of time, and decreasing a value of theintegrator over a second period of time; and in the second mode,providing an output of a second dynamic controller to the actuator viathe second interface of the processor, wherein the output of the seconddynamic controller is proportional to the received pilot control signal.15. The flight control system of claim 14, wherein; the first value is atranslational speed; the second value is an attitude; and the firstcondition of the rotorcraft is a velocity of the rotorcraft.
 16. Theflight control system of claim 15, wherein: the attitude comprises apitch attitude and a roll attitude; executing the first dynamiccontroller comprises executing a first pitch dynamic controller having apitch integrator and executing a first roll dynamic controller having aroll integrator; and providing the output of the first dynamiccontroller to the actuator via a second interface of the processorcomprises providing an output of the first pitch dynamic controller to apitch actuator and providing an output of the first roll dynamiccontroller to a roll actuator.
 17. The flight control system of claim14, wherein the executable program further includes instructions to fadein a gain of the pilot control signal in the second dynamic controllerwhen a pilot control stick is out of detent; and fade out a gain of thepilot control signal in the second dynamic controller when the pilotcontrol stick is in detent.
 18. The flight control system of claim 14,wherein the executable program further includes instructions totransition from the second mode to the first mode when the firstcondition of the rotorcraft decreases below a second predeterminedthreshold, wherein transitioning from the second mode to the first modecomprises: fading in the gain of the first dynamic controller over athird period of time, and reactivating the integrator.
 19. The flightcontrol system of claim 18, wherein the first predetermined threshold isgreater than the second predetermined threshold.
 20. The flight controlsystem of claim 14, wherein the first dynamic controller furthercomprises a proportional path, and an output of the first dynamiccontroller is a sum of the proportional path and the output of theintegrator.
 21. The flight control system of claim 14, whereindecreasing the value of the integrator comprises coupling the output ofthe integrator to an input of the integrator via a first feedback path.22. The flight control system of claim 21, wherein the first feedbackpath comprises a limiter.
 23. The flight control system of claim 14,wherein the first predetermined threshold comprises a ground speed. 24.The flight control system of claim 14, wherein the first period of timeand the second period of time are between 1 s and 10 s.
 25. A rotorcraftcomprising: a body; a power train coupled to the body and comprising apower source and a drive shaft coupled to the power source; a rotorsystem coupled to the power train and comprising a plurality of rotorblades; a flight control system operable to change at least oneoperating condition of the rotor system; a pilot control assemblyconfigured to receive commands from a pilot, wherein the flight controlsystem is a fly-by-wire flight control system in electricalcommunication with the pilot control assembly; and a flight controlcomputer in electrical communication between the flight control systemand the pilot control assembly, the flight control computer configuredto: receive, from the pilot control assembly a pilot command to change afirst flight characteristic, when a velocity of the rotorcraft is lessthan a first velocity threshold, interpret the first flightcharacteristic as a translational speed in a first mode and determine anattitude based on the translational speed via using an attitudecontroller comprising an integrator and a proportional path parallel tothe integrator, when the velocity of the rotorcraft is greater than asecond velocity threshold, interpret the first flight characteristic asa rate in a second mode, and when the velocity of the rotorcraftincreases past the second velocity threshold, fade out a gain of theattitude controller, and successively decrease a value of theintegrator.
 26. The rotorcraft of claim 25, wherein: the attitudecomprises a pitch attitude; and the rate comprises a pitch rate.
 27. Therotorcraft of claim 26, wherein: the attitude further comprises a rollattitude; and the rate further comprises a roll rate.
 28. The rotorcraftof claim 25, wherein the flight control computer is further configuredto transition from the second mode to the first mode when the velocityof the rotorcraft decreases below the first velocity threshold, andtransitioning from the second mode to the first mode comprises: fadingin the gain of the attitude controller, and reactivating the integrator.29. The rotorcraft of claim 28, wherein the first velocity threshold isless than the second velocity threshold.
 30. The rotorcraft of claim 25,wherein the flight control computer is further configured to: fade in again of the pilot command to an attitude rate controller when a pilotstick of the pilot control assembly is out of detent; and fade out again of the pilot command to the attitude rate controller when the pilotstick of the pilot control assembly is in detent.